The aspects of the disclosed embodiments relate to a method for manufacturing an aircraft fuselage section of composite material that permits the introduction of systems and elements of the aircraft and/or robots capable of installing these systems and elements into the interior of the section.
The disclosed embodiments find applications in the field of aeronautics and in particular, in the field of designing aircraft fuselage sections of composite material.
An aircraft fuselage is traditionally a tubular body with a skin fastened to an internal structure. The skin is made of metal panels mounted and fastened around the metal internal structure, called the internal skeleton of the aircraft. The various panels are made singly and then assembled with one another by an overlapping process to constitute a fuselage section. Several sections are assembled with one another to form the entire fuselage. Such a metal fuselage has the drawback of being heavy since it is wholly metallic. It also has the drawback of requiring joints between the metal panels and between the sections, which creates overlapping areas with excess thickness, further increasing the weight of the aircraft.
To reduce the weight of the fuselage, aerodynamic builders try to replace certain metal elements by elements made of composite material. Composite materials can be used to make one or more parts of an aircraft fuselage. These parts of the fuselage, called sections, are made from dry fiber sheets pre-impregnated with thermosetting resin. These fiber sheets are placed in molds and then heated. Under the action of heat, the resin polymerizes, which permits the fiber reinforcement to retain the shape of the mold.
Two methods are known at this time for making aircraft fuselage sections of composite material. A first method consists of making panels of composite material singly, which are then assembled by a technique of overlapping assembly, essentially similar to that described above for metal panels. FIG. 1 represents an example of an aircraft made from panels. In this figure, each Pnumber reference corresponds to a fuselage panel. In this example, the sections T1, T2, and T3 are each made from four panels.
A second method consists of making the sections of composite material from a single piece. Actually, the techniques for making the elements of composite material permit making parts with large dimensions and complex shapes. Accordingly, it is possible to manufacture a tapered portion of fuselage, or section. In this case, each section is made of a single part in a single step. These sections are called one-shot or single-piece sections. Such a section is manufactured from fiber sheets pre-impregnated with resin, wound around a mold that has the desired tapered shape. The mold can be a male mold with the shape of a tubular cylinder, for example. When the fiber sheets have been wound around the mold, they are heated and then cooled. After cooling, the skin of composite material is detached from the mold, either by sliding it off or by disassembling the mold.
In this method, each section is made as a unit. Several sections are then assembled with one another to make up the fuselage.
This method has the advantage of eliminating the phase of assembly by overlapping panels, which avoids the use of joining elements and the creation of excess thicknesses, thus reducing the total weight of the fuselage. It also permits having continuous fibers that are not cut off and that consequently show improved performance. It also has the advantage of not requiring any overlap between the longitudinal joints and the circumferential joints, which is often disadvantageous in terms of weight.
This method of manufacturing a one-piece section also has the advantage of being fast. Actually, winding on a male mold is fast, and the number of assembly operations, costly in terms of manufacturing steps, is reduced (no longitudinal joints).
This technology for manufacturing one-piece sections accordingly provides many advantages, in particular when a large number of fuselages are to be produced. However, it also has drawbacks. A first drawback relates to the assembly of two one-piece sections, which requires stringent manufacturing tolerances that are difficult to respect. More precisely, the perimeter of the two sections to be assembled has to rigorously identical; a defect of shape can be corrected by the flexibility of the composite material, but a difference in the perimeter can absolutely not be corrected. Actually, in contrast to metal panels, it is not possible in this procedure to disconnect certain panels, in other words to undo the joint of the panels, to flex the assembly, and compensate for any differences from the manufacture of each section.
Another drawback of this procedure concerns the installation of interior equipment (struts, floors, etc.) and internal systems (electric, hydraulic, or pneumatic systems) inside the section. Actually, all of these systems and equipment are installed in the interior of the section after the section is manufactured, via the lateral openings of said section, i.e. the two openings located at one end and the other of the section. The parts to be installed, installation robots, and even personnel have to enter the section through the lateral openings.
An example of a one-piece section is shown in FIG. 2. In this example, there is the skin 1 of the section and a lateral opening 3. Installation equipment and elements 2 are also shown, introduced into the skin 1 through the lateral opening 3. As shown in this figure, the lateral openings correspond to a closed circular cross section. The installation of equipment and systems accordingly proves to be difficult, in particular for voluminous elements.